1. Field of the Invention
This invention relates to coolable hollow gas turbine engine turbine blades and, more particularly, to pins extending between pressure and suction side walls of the blades.
2. Description of Related Art
Gas turbine engines typically employ a row of coolable hollow turbine blades secured to an outer perimeter of a rotor disk with a stationary turbine nozzle having a plurality of stator vanes disposed upstream therefrom. The combustion gases flow between the stator vanes and between the turbine blades for extracting energy to rotate the rotor disk. The temperatures within gas turbines may exceed 2500 degrees Fahrenheit and cooling of turbine blades is very important in terms of blade longevity. Without cooling, turbine blades would rapidly deteriorate. Cooling for turbine blades is very desirable and much effort has been devoted by those skilled in the blade cooling arts to devise geometries for internal cavities within turbine blades in order to enhance cooling. Since the combustion gases are hot, the turbine vanes and blades are typically cooled with a portion of compressor air bled from the compressor for this purpose. Diverting any portion of the compressor air from use in the combustor necessarily decreases the overall efficiency of the engine. It is desirable to cool the vanes and blades with as little compressor bleed air as possible.
Typical turbine vanes and blades include an airfoil over which the combustion gases flow. The airfoil, typically, includes one or more straight through channels and serpentine cooling passages through which cooling air from compressor bleed air is channeled for cooling the airfoil. The airfoil may include various turbulators therein for enhancing cooling effectiveness and the cooling air is discharged from the passages through various film cooling holes disposed around the outer surface of the airfoil. In pursuit of higher cooling effectiveness, modern blades have led to multi-pass cooling circuits.
It is also known to pass the cooling air through serpentine cooling air circuits and other passages in the interior of the blade which warms up the cooling air as it travels through the passages before being impinged on the leading edge of the blade. The temperature difference across the leading edge is lower than directing cooling air through the root of the blade for impingement resulting in lower thermal stresses in the blade leading edge and the life of the blade is enhanced. This makes efficient use of cooling flow since the flow is able to internally cool the blade over much of the blade mid-span before flowing out radial leading edge cooling holes to film cool the blade airfoil externally.
Known turbine airfoil cooling techniques include the use of internal cavities forming a serpentine cooling circuit. Particularly, serpentine passages, leading edge impingement bridges, turbulence promoters and turbulators, film holes, pins, and trailing edge holes or pressure side bleed slots are utilized for blade cooling.
The hollow turbine blades are subject to resonance at natural frequencies of the blade and it is known to modify blades and designs thereof to avoid operating at the natural frequencies of the blade for other than transient periods during engine operation. It is desirable to cool turbine blades with as little cooling air as possible and which substantially avoids operating at the natural frequencies of the blade for other than transient periods during engine operation.